1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a seal for a rotor disk in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine engine, includes a compressor to deliver compressed air to a combustor that produces a hot gas flow that is then passed through a turbine to produce mechanical power. The turbine includes a number of stages or rows of rotor blades and stator vanes that form a hot gas path through the turbine. The rotor blades form a seal with a stationary part of the engine to limit egress of the hot gas flow into parts of the engine that can be thermally damaged.
One prior art seal used in a gas turbine engine is where the rotor disk includes a labyrinth seal having a number of knife edges that rotates near to a surface on the stationary casing to form a rotary seal. The knife edge seal limits the leakage of flow but does not totally block the leakage. Brush seals are also used to reduce leakage. However, brush seals make contact with the rotating part and therefore cause wear of the brush bristles. Also, brush seals do not make good seals at high rotational speeds. One major problem with this type of rotary seal used in a gas turbine engine is that the gap formed between the rotary seal can vary depending upon the engine temperatures. During engine transients, the knife edges can actually rub against the stationary seal interface and thus cause heating or damaged to the knife edges. Some complex arrangement of parts have been proposed in the prior art to limit the seal gap in these types of rotary seals in gas turbine engines.